US9140196B2ActiveUtilityA1
Method for controlling a gas turbine and gas turbine for implementing the method
Est. expiryOct 27, 2030(~4.3 yrs left)· nominal 20-yr term from priority
Inventors:Herbert Fuchs
F02C 9/28F05D 2270/303F05D 2220/3212F05D 2270/3015
54
PatentIndex Score
1
Cited by
15
References
25
Claims
Abstract
A method for controlling a gas turbine, including during transient operating states, and such a gas turbine are provided. The gas turbine includes a compressor for compressing inducted combustion air, a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air, and a multistage turbine for expanding the generated hot gas and performing work. The controlling of the gas turbine is carried out in accordance with the hot gas temperature which is derived from a plurality of other measured operating variables of the gas turbine. A reliable controlling of the gas turbine is achieved, even during rapid changes, by pressure measurements being gathered exclusively at different points of the gas turbine for derivation of the hot gas temperature.
Claims
exact text as granted — not AI-modifiedWhat is claimed is:
1. A method for controlling a gas turbine, wherein the gas turbine comprises a compressor for compressing inducted combustion air, a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air, and a multistage turbine for expanding the generated hot gas and performing work, wherein the method comprises:
controlling the gas turbine in accordance with a hot gas temperature which is derived from a plurality of other measured operating variables of the gas turbine,
wherein a derivation of the hot gas temperature is formed exclusively through pressure measurements at different spatial points in a flow direction in a region of hot gas flow of the gas turbine,
wherein the pressure measurements are carried out in the hot gas flow which is generated in the combustion chamber, and
wherein the pressure measurements include a pressure loss measurement in a main flow of the gas turbine.
2. The method as claimed in claim 1 , wherein the measuring positions, between which the pressure loss is determined, lie downstream of flames which are present in the combustion chamber of the gas turbine.
3. The method as claimed in claim 2 , wherein the measuring positions, between which the pressure loss is determined, lie exclusively in the combustion chamber.
4. The method as claimed in claim 3 , wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured.
5. The method as claimed in claim 4 , wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula
T
=
const
**
(
n
)
·
(
p
Δ
p
)
2
·
(
p
·
p
amb
0
p
0
·
p
amb
)
1
,
31
-
1
1
,
31
=
F
(
p
,
Δ
p
)
wherein n is the rotational speed, p amb is the ambient pressure and p 0 is the pressure upstream of the turbine and p amb0 is the associated ambient pressure, both at a nominal point.
6. The method as claimed in claim 2 , wherein a turbine flange is located in the gas turbine at the transition between the combustion chamber and the turbine, and
wherein the measuring positions, between which the pressure loss is determined, lie exclusively in the turbine flange.
7. The method as claimed in claim 6 , wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured.
8. The method as claimed in claim 7 , wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula
T
=
const
**
(
n
)
·
(
p
Δ
p
)
2
·
(
p
·
p
amb
0
p
0
·
p
amb
)
1
,
31
-
1
1
,
31
=
F
(
p
,
Δ
p
)
wherein n is the rotational speed, p amb is the ambient pressure and p 0 is the pressure upstream of the turbine and p amb0 is the associated ambient pressure, both at a nominal point.
9. The method as claimed in claim 2 , wherein a turbine flange is located in the gas turbine at the transition between the combustion chamber and the turbine, and
wherein the measuring positions, between which the pressure loss is determined, lie both in the combustion chamber and in the turbine flange.
10. The method as claimed in claim 9 , wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured.
11. The method as claimed in claim 10 , wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula
T
=
const
**
(
n
)
·
(
p
Δ
p
)
2
·
(
p
·
p
amb
0
p
0
·
p
amb
)
1
,
31
-
1
1
,
31
=
F
(
p
,
Δ
p
)
wherein n is the rotational speed, p amb is the ambient pressure and p 0 is the pressure upstream of the turbine and p amb0 is the associated ambient pressure, both at a nominal point.
12. The method as claimed in claim 2 , wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured.
13. The method as claimed in claim 12 , wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula
T
=
const
**
(
n
)
·
(
p
Δ
p
)
2
·
(
p
·
p
amb
0
p
0
·
p
amb
)
1
,
31
-
1
1
,
31
=
F
(
p
,
Δ
p
)
wherein n is the rotational speed, p amb is the ambient pressure and p 0 is the pressure upstream of the turbine and p amb0 is the associated ambient pressure, both at a nominal point.
14. The method as claimed in claim 1 , wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured.
15. The method as claimed in claim 14 , wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula
T
=
const
**
(
n
)
·
(
p
Δ
p
)
2
·
(
p
·
p
amb
0
p
0
·
p
amb
)
1
,
31
-
1
1
,
31
=
F
(
p
,
Δ
p
)
wherein n is the rotational speed, p amb is the ambient pressure and p 0 is the pressure upstream of the turbine and p amb0 is the associated ambient pressure, both at a nominal point.
16. The method as claimed in claim 1 , wherein in the pressure measurements, possible pulsation signals, which emanate from combustion chamber pulsations, are suppressed by means of one of time-based averaging and filtering.
17. The method as claimed in claim 1 , wherein the controlling of the gas turbine comprises controlling the gas turbine during transient operating states of the gas turbine.
18. The method as claimed in claim 1 , wherein the different spatial points are arranged in series in the flow direction.
19. The method as claimed in claim 1 , wherein the pressure measurements are made through at least two different spatial points in a flow direction in a region of hot gas flow.
20. A gas turbine comprising:
a compressor for compressing inducted combustion air;
a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air;
a multistage turbine for expanding the generated hot gas and performing work;
a machine control unit;
at least two pressure sensors for measuring a pressure loss in the hot gas flow, the pressure sensors being arranged at different points in series in a flow direction in the region of the hot gas flow which is generated in the combustion chamber, and the pressure sensors being connected to the machine control unit; and
a turbine flange arranged in the gas turbine at a transition between the combustion chamber and the turbine,
wherein the machine control unit is configured to control the gas turbine in accordance with the hot gas temperature which is derived from a plurality of other measured variables of the gas turbine,
wherein a derivation of the hot gas temperature is formed exclusively by means of pressure measurements at the different points of the gas turbine, and
wherein at least one of the pressure sensors is located in a region of the turbine flange.
21. The gas turbine as claimed in claim 20 , wherein the pressure sensors are located in the region of the combustion chamber.
22. The gas turbine as claimed in claim 20 , wherein the pressure sensors are located in the region of the turbine flange.
23. The gas turbine as claimed in claim 20 , comprising:
a third pressure sensor located directly upstream of a first blade row of the turbine and being connected to the machine control unit of the gas turbine.
24. The gas turbine as claimed in claim 23 , wherein the pressure sensors are connected to the machine control unit of the gas turbine via a device for suppressing pulsation signals.
25. The gas turbine as claimed in claim 20 , wherein the pressure sensors are connected to the machine control unit of the gas turbine via a device for suppressing pulsation signals.Cited by (0)
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